Device for sonic boom reduction and improving aircraft performance

ABSTRACT

Method and means for improving the performance of, and particularly for reducing the sonic boom produced by a supersonic aircraft, comprising the production of a jet stream of approximately equal pressure but higher Mach number than the ambient supersonic flow, which steam is directed below the wing leading edge to intercept and interact with the wing shock wave. The interaction weakens the wing shock and decreases its propagation velocity, so that the wing shock, which normally reinforces the leading nose shock at the ground, will be shifted aftward spatially and in time in the boom signature, and will be delayed in its arrival with respect to the leading nose shock at any fixed position on the ground. By thus altering the signature, both the peak overpressure and the positive impulse of the boom may be substantially reduced without affecting the lift in support of the aircraft. Indeed, proper use of the stream will increase the lift on the aircraft, permitting some decrease in the angle of attack in maintaining level flight, which further weakens the wing shock to add to the altering of the boom signature. Also, the increase in lift, occurring largely on the aft part of the wing, results in reduction in drag, aftward shift of the center of pressure, and increase in the aerodynamic efficiency factor or lift-to-drag ratio, so that the jet stream can be used to improve various aspects of aircraft performance. Supersonic aircraft equipped with the anti-boom means, and a particular feasible means for producing the antiboom jet using existing technology are described, along with particular alternatives for the manner of incorporating the means on existing supersonic transports.

United States Patent [191 Cheng 51 June 5, 1973 [54] DEVICE FOR SONIC BOOM REDUCTION AND IMPROVING AIRCRAFT PERFORMANCE [75] Inventor: Sin-I Cheng, Princeton, NJ.

[73] Assignee: Research Corporation, New York,

[22] Filed: June 15, 1970 a [211 App]. No.2 46,293

[52] U.S. Cl. ..244/1 N, 244/55 [51] Int. Cl ..B64c 21/00 [58] Field of Search ..244/l N, 53 R, 54,

244/55; 181/33 H, 33 HB, 33 HC [56] References Cited UNITED STATES PATENTS 3,510,095 5/1970 Chuan ..244/l30 3,596,852 8/1971 Wakefield ..244/l N Primary ExaminerDuane E. Reger Assistant Examiner-Barry L. Kelmachter Att0mey1-Ienry T. Burke, Robert S. Dunham, P. E. l-lenninger, Lester W. Clark, Gerald W. Griffin, Thomas F. Moran, Howard J. Churchill, R. Bradlee Boal, Christopher C. Dunham and Thomas P. Dowd [57] ABSTRACT Method and means for improving the performance of, and particularly for reducing the sonic boom DUAL EA/G/A/E UA//7; MKL. AMT/500M JET MEANS produced by a supersonic aircraft, comprising the production of a jet stream of approximately equal pressure but higher Mach number than the ambient supersonic flow, which steam is directed below the wing leading edge to intercept and interact with the wing shock wave. The interaction weakens the wing shock and decreases its propagation velocity, so that the wing shock, which normally reinforces the leading nose shock at the ground, will be shifted aftward spatially and in time in the boom signature, and will be delayed in its arrival with respect to the leading nose shock at any fixed position on the ground. By thus altering the signature, both the peak overpressure and the positive impulse of the boom may be substantially reduced without affecting the lift in support of the aircraft. Indeed, proper use of the stream will increase the lift on the aircraft, permitting some decrease in the angle of attack in maintaining level flight, which further weakens the wing shock to add to "the altering of the boom signature. Also, the increase in lift, occurring largely on the aft part of the wing, results in reduction in drag, aftward shift of the center of pressure, and increase in the aerodynamic efficiency factor or lift-to-drag ratio, so that the jet stream can be used to improve various aspects of aircraft performance. Supersonic aircraft equipped with the antiboom means, and a particular feasible means for producing the antiboom jet using existing technology are described, along with particular alternatives for the manner of incorporating the means on existing supersonic transports.

50 Claims, 36 Drawing Figures PATENTEDJUN sms 3.731.119

' sum 03 or 10 PATENTEDJUH 5 I975 3,737,119

sum as or 10 PATENTEDJUN 5 I975 ANTI-500M SHEET 07 [1F 10 Q5 ENG/NE JET ME 4M c 41 FREE STREAM Q PATENTEUJUN 5 I975 saw 0a or 10 PATENTEBJUH 5 197a 3.737.119

' sum 10 or 10 DEVICE FOR SONIC BOOM REDUCTION AND IMPROVING AIRCRAFT PERFORMANCE BACKGROUND OF THE INVENTION The present invention relates to supersonic aircraft and more particularly to a method and means for reducing the sonic boom associated therewith, and for improving the flight characteristics.

With the advent of supersonic flight, to the many conventional problems of aircraft designers, such as structural strength, lift-to-drag considerations, etc., was added the problem of sonic boom. Sonic boom refers to the pressure disturbances created at large distances from an aircraft by the shock waves associated with the aircraft in supersonic flight. More particularly, when an aircraft flies supersonically through the quiescent air, the air molecules do not receive any signal or warning to'yield the space they occupy to the oncoming aircraft. The air molecules struck by the aircraft are suddenly displaced, piled up on the leading surfaces and then forced to move over the surface contours of the aircraft. The sudden piling up of these air molecules creates regions of sudden increase of air density, temperature and pressure about the aircraft, which regions are outwardly bounded by what are technically called shockwaves. The highly compressed air in these regions tends to expand and compress the surrounding air resulting, in time, in the propagation of the shock waves into the surrounding quiescent atmosphere at a speed somewhat greater than the speed of sound but still lagging behind the aircraft which created them. Also, as the aircraft passes, the compressed air over-expands along any decreasing transverse sections of the various elements of the aircraft surface and the over-expanded air from the top and the bottom surfaces of each element impinge upon each other, producing trailing shocks in the process of becoming realigned. This multitude of shock waves propagates away from the passing aircraft while new ones are continuously generated from the various points of the leading and the trailing surfaces as the aircraft proceeds. The resulting complex of shock waves appears as a series of sheets having curved surfaces emanating from the points on the aircraft from which they are generated. At sufficiently large distances from the aircraft these shock surfaces tend to merge and appear conical. The surfaces of the cones are more or less oriented in the Mach direction of the local supersonic stream, that is, in the direction of a wave traveling at Mach one in the stream. For this reason, such cones are often referred to as Mach cones, although somewhat incorrectly since as previously noted, they propagate at a speed somewhat greater than the speed of sound. The portions of these Mach cones beneath the aircraft may eventually strike the ground, if not previously dissipated or merged in the atmosphere, producing a sonic boom. The sonic boom is therefore a composite of pressure waves caused by the various components of the aircraft which generate many elementary wave systems whose pressure profiles are in the form of N-waves, that is, the temporal or the spatial sequence of pressure variation approaches the shape of the letter N. Each N-wave is composed of a forward region of overpressure, that is, above the ambient, due to the sudden pressure rise across a leading shock, followed by a gradual fall generally to below the ambient through a spread out expansion wave behind the shock, and often terminated by a sudden rise across a trailing shock. These elementary N-waves are of different strengths, scales and spacings superposed one on top of the other. Each elementary wave system decays in strength while propagating away from the aircraft by engulfing and energizing more and more air. The waves interact and merge. A notable nonlinear feature of the propagation of these wave systems is that various leading shocks, if not annihilated in the merging process, will eventually catch up with their preceding shocks and reinforce one another to form a single leading nose shock for the entire system in a homogeneous atmosphere. Thus, at a distance, say 500 aircraft lengths below the aircraft, a single N-wave emerges as an asymptotic wave form in the so-called far field. Of course, at distances of one to two hundred aircraft lengths for example, some of the major shocks may not have coalesced so that in this region the pressure variation may still consist of several individual shocks. This is the mid field characteristic of the pressure variation. The complex shock pattern in the immediate vicinity of the aircraft is referred to as its near field characteristic. I

A detailed description of the pressure variation, either spatially or temporally, is referred to as the boom signature. The largest overpressure reached above the ambient is called the peak overpressure and the time required for the signature to reach the peak overpressure is called the rise time. The integral of the force during the overpressure period is called the positive impulse, corresponding to the positive area under the signature. These characteristics which make .up the sonic boom from any aircraft, will clearly depend on the aircraft design configuration and the flow field around it, as well as the atmospheric conditions through which the wave system must pass before reaching an observer on the ground.

The intensity of sonic booms produced by some supersonic military aircraft flying at some 30,000 to 40,000 feet, have been found to be in the range of 2.0 psf (pounds per square foot) of overpressure. Overpressures of this magnitude and above, have generally been found in flight tests to be objectionable. T o significantly reduce the overpressure, such supersonic military aircraft can fly at much greater altitudes. However, for much heavier supersonic transports which are already designed for cruising at high altitudes, in the range of from 60,000-to 70,000 feet, a problem arises.

For the commercial supersonic transports currently under development, the peak overpressure during cruise is expected to be generally less than 2 to 3 psf at ground level, well below values that might possibly cause physical damage to structures or even to sensitive organs of animals or people. Booms of the range, 2 to 3 psf, are basically an annoyance, particularly due to the element of surprise. The extent of annoyance varies with environments and subjective judgments, so that an annoyance index is exceedingly difficult to define. For example, the loudness of a sonic boom perceived by an observer apparently depends upon the initial rise time of the overpressure regions of its signature, since signatures with shorter rise times sound louder than those with the same peak overpressure and duration, but with longer rise times. This may be due to the fact that sonic booms with shorter rise times may have more energy in the frequency bands to which the human ear is most sensitive and therefore many'people may regard such booms as highly annoying. 0n the other hand, the

longer rise time signatures may be richer in frequencies at the lower end of thefrequency spectrum, and beyond, and although hardly audible, their vibratory inputs may be as annoying or even more so to others. The difference in reactions will depend on the physical sensitivities of various observers. Their psychological dispositions at a given time, further complicate the matter. Whether some forewarning of the boom, which will put an observer under anticipation, is favorable or unfavorable is uncertain and dependent again on his subjective judgment. The environment will not only influence the psychological disposition of an observer, but also modify the nature of the boom signature. After traversing a structure such as a building, the relative constituency of various frequency components will be quite distorted and the dominant component inside the structure may be much different from that outside. Thus, subjects indoors and outdoors will have different responses to the same boom signature. Also, not to be overlooked is the fact based on laboratory tests, that repeated exposure to boom appears to increase the annoyance threshold and the tolerable levels of human subjects. The delineation of acceptable versus objectionable booms therefore, is so complex that after considerable detailed analysis, neither the scientific community nor the appropriate regulatory agencies can identify the most objectionable characteristics which should be reduced in preference to others.

Based on the latest published information, however, the following observations may serve somewhat as ground rules for interpreting the meaning of reduction of sonic boom. It has been definitely determined that the acceptability of the boom is generally increased by: l) reduction of peak overpressure (or underpressure if larger); and (2) reduction of positive impulse. Also, among boom signatures of similar peak pressure levels, detailed features such as the absence of spikes, the presence of finite rise time (10 millisec. or larger), more rounded and stretched out signatures or shockless signatures are often judged more acceptable. However, although some believe that shockwaves should be eliminated even at the expense of higher positive impulse, all of these detailed features of boom signatures do not seem nearly so important as the two major gross characteristics cited. Further, these detailed features will be seriously distorted, if not completely destroyed, upon traversing an intervening structure, and will not be of much significance in causing annoyance to subjects indoors. To subjects indoors, the positive impulse, being a temporarily integrated property of the boom signature, will be of concern and may indeed be more meaningful as an index of annoyance than peak overpressure, although the two are generally related at least for roughly similar signatures. A peak over pressure level of 1.5 to 1.7 psf has been correlated in laboratory tests and cited as equivalent to 1 l PNDB (perceived noise in decibels) which is widely regarded as the maximum tolerable noise level. Thus, 1.5 psf has been suggested and quoted as the maximum tolerable boom overpressure level, but it will be seen that such an arbitrary gross characteristic is clearly not a satisfactory standard in view of the various factors previously discussed.

It is well established, however, that boom signatures with overpressure levels less than l psf are really not much of a nuisance especially for subjects indoors. Therefore, the reduction of boom overpressures for commercial supersonic transports to below 1.5 psf should be considered as the minimum objective and possibly to l psf as the ideal objective of any boom reduction devices if accomplished by the reduction of positive impulse to similar amounts.

In any event, it has been found that the aircraft design factors that particularly affect the intensity or severity of a sonic boom are the length and weight of the aircraft, as well as its speed and altitude. For the asymptotic N-wave form in the far field, that is, in the vicinity of the ground, theoretical and experimental studies have led to a reasonable estimate of the peak overpressure and of its qualitative dependence on various parameters. The peak overpressure may be roughly divided into two interrelated parts:

a. A p which is the component due to the airplane volume at zero lift; and

b A p which is the component due to the angle of attack of the airplane in producing lift.

In a real atmosphere, estimates based on such theoretical studies show that while for an aircraft flying at low altitudes, A p dominates, at the high cruising altitudes of commercial supersonic transports (approximately 60,000 feet) and for such heavy aircraft, A 2, becomes the larger component. Thus, for the reduction of the boom from a cruising commercial supersonic transport, at ground level, the overpressure component due to the lift, that is, the lift boom," A p,,, is of particular concern. However, previous attempts at devising means for reducing the sonic boom have avoided operating on this component directly for fear of destroying the necessary lift on the aircraft.

Although most earlier proposals to reduce sonic booms have been based on theoretical estimates of the asymptotic overpressure, foremost among which are the concepts of early warning and of phantom planes, they have been directed to displacing and/or weakening the leading nose shock and the tail or trailing shocks behind the aircraft while leaving the dominant wing shock unchanged. Various modes of implementing such concepts through conventional and exotic schemes have been proposed. However, none of these schemes have been favorably'considered by the manufacturers, since none has met the two essential criteria of a practical solution, that is, that it be scientifically sound and be capable of feasible incorporation in a supersonic aircraft without upsetting the general performance requirements of the aircraft. The performance requirements of a successful commercial supersonic transport, for example, are highly restrictive so that the permissible variations in design parameters are quite marginal. Very little penalty in the form of weight or drag increase can be tolerated for boom reduction purposes. Hence, while another prior proposal, the multi-wing concept, is attractive and promises a meaningful reduction of the boom, it is not a practical solution because of the weight penalty of the complicated wing structure and/or the drag increase at off-design operation.

In view of the fact that the lift boom, A p,,, is the dominant component of the boom from a cruising supersonic transport, then the present invention accordingly embodies a method for dealing with the lift boom to reduce the peak overpressure and the positive impulse, which method will be shown to be conceptually sound. In addition, a preferred means for accomplishing this method is presented which may be incorporated into the currently designed supersonic transports, or SSTs, without further upsetting their economic desirability. The approach taken is directed toward the weakening of the wing shock so as to decrease its propagation velocity while avoiding such interactions with the shockwaves as would cause a loss of the lift which must be derived from the wing to keep the aircraft in level flight. In fact, the present invention may provide increased lift, and improve the lift-to-drag ratio and the overall performance of the aircraft in the process to at least compensate for the unavoidable weight penalty caused by the incorporation of the boom reduction means on the aircraft.

SUMMARY OF THE INVENTION The present invention involves the production of a jet stream of approximately equal pressure but higher Mach number than the ambient supersonic flow about the jet and the directing of this high speed jet stream toward the wing shock wave below the wing leading edge. The high speed jet stream, which will be called the anti-boom jet, intercepts the wing shock wave and produces an intense interaction. The interaction creates a pressure profile on the underside of the wing which maintains and, in fact, may, under proper conditions, improve the lift. The shock wave that emerges from the high speed jet stream is a refracted continuation of the leading wing shock, and is weaker than the wing shock without interaction, i. e., in the absence of the anti-boom jet. This weakened wing shock will propagate into the ambient atmosphere at a lesser velocity than the stronger wing shock, so that the boom signature at the ground level will be altered, since the wing shock wave, which normally reinforces the leading nose shock wave, will now be shifted aftward and separated from the leading nose shock. The slightly weakened and shifted wing shock will rise from a lower pressure level in the boom signature and reach a maximum overpressure at a level less than that resulting when the wing shock and nose shock are superposed. The maximum underpressure in the rear half of the boom signature can likewise be reduced and the boom signature will be stretched out downstream. With this method, the maximum overpressure caused by cruising commercial supersonic transports may be reduced appreciably below the 1.5 psf level and thus well into an humanly acceptable range.

The means for producing the high Mach number jet stream is located in whole or in part upstream of the wing shock wave and directs the anti-boom jet to interact with the shock wave emanating from the leading edge of the wing at appropriate distances therebelow. Various embodiments are possible and may comprise a single high pressure source in the fore section of the fuselage or may include in part some high pressure source carried elsewhere on board the aircraft. Since a major portion of the fluid flux in the anti-boom jet can be air economically captured from the ambient atmosphere, the main thrust-producing engines of the aircraft, with power capacity available may contain all the elements required to produce the jet and can be conveniently used for the purpose if properly located. If the excess power capacity is sufficient, no major hardware will be required which would result in a dead weight penalty. The energy may be derived from the same fuel as is required for the regular operation of the aircraft and the additional thrust derived from the anti-boom jet permits reduction of the thrust required to be derived from the regular operation of the jet engines. While,-thermodynamically, the cruising thrust may be derived at a lower efficiency or a higher specific fuel consumption than the optimum choice of the aircraft designer without boom reduction consideration, the additional fuel expenditure in operating the anti-boom jet at the same total thrust level will be the expendable item in the price paid for boom reduction. Presently, the engines for the current commercial supersonic transports are all turbojets with afte'rburners which are adaptable to operation with the anti-boom jet means. However, if the anti-boom jet operation is to be incorporated into the main engines, other composite power plants, such as turbofan engines, may prove to be more advantageous overall, depending very much on the specific aircraft. Also, the engines exhaust may be used with the anti-boom jet in weakening the wing shock wave.

In any event, a preferred embodiment is presented indicating the manner in which the new General Electric Company prototype GE4/J5P engine may be adapted to produce an anti-boom jet for satisfactory use on the prototype 2707-300 supersonic transport presently being manufactured by the Boeing Company.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 illustrates an aircraft flying at supersonic speed and shows the propagation of the major shock waves to the ground level;

FIG. 2 is a plot of the contribution made to the overpressure in a boom signature due to the volume of the aircraft and the lift thereon as a function of altitude;

FIGS. 3 a-e and FIGS. 4 a-e are schematic diagrams of boom signatures illustrating the effect on the peak over-pressure and the positive impulse for various shifts of the wing boom with respect to the nose boom;

FIG. 5 is a diagram illustrating the propagation through the atmosphere of a nose shock front and a wing shock front from an aircraft with respect to each other and their associated Mach cones;

FIG. 6 is a diagram illustrating the components of propagation of a Mach cone and a shock wave in the atmosphere;

FIG. 7a is a representative two-dimensional section across a bi-convex wing of an aircraft in supersonic cruise at an angle of attack (1 illustrating the resulting shock waves and expansion waves forward and aft of the dividing Mach line of the wing;

FIG. 7b is a plot illustrating the surface pressure over the upper and lower surfaces of the wing shownin FIG.

FIG. 7c is aplot illustrating the resultant lift force due to the pressure difference on the upper and lower surfaces of the wing shown in FIG. 7a and also illustrating the additional lift to be derived in accordance with the present invention;

FIG. 7d is a plot as in FIG. 7c, illustrating the lift force distribution over thewing for a reduced angle of attack in accordance with the present invention;

FIG. 8a is an illustration of the interaction between a wing shock wave and a contact discontinuity produced by a stream having a lower Mach number than that of its ambient flow past a wing of an aircraft in supersonic cruise.

FIG. 8b is a plot known as a p-& shock polar illustrating the pressure ratio across the shock wave shown in FIG. 8a as a function of the angle of deflection of the flow for different Mach numbers;

FIG. 86 is an illustration as in FIG. 8a, wherein the contact discontinuity is produced by a stream of higher Mach number than that of the ambient flow past the wing of the aircraft;

FIG. 8d is a p-o shock polar for the interaction illustrated in FIG. 80;

FIG. 9a is an illustration of the interaction of a shock wave in a supersonic stream passing through a series of contact discontinuities produced by successive streams having a higher Mach number and a lower Mach number respectively than that of the supersonic stream;

FIG. 9b is a p-S shock polar for the interactions illustrated in FIG. 90;

FIG. 10a is a more detailed illustration of the interaction of a shock wave from a bi-convex airfoil with a contact discontinuity produced by a stream of higher Mach number than that of the flow stream passing by the airfoil;

FIG. 10b is a p-o shock polar for the interaction illustrated in FIG. 10a;

FIG. 11 is a diagrammatic view of anti-boom jetproducing system utilizing the main thrust-producing engines mounted forward on the aircraft and as incorporated on one side of a supersonic transport of the form of the Boeing Company prototype 2707-300;

FIG. 12 is a partial side view taken along the lines l2--l2 in FIG. 11, showing the interaction of the antiboom jet and the engine exhaust with the shock wave from the wing leading edge;

FIG. 13 is a sectional view taken along the lines 13-13 in FIG. 11, illustrating the relationship of the anti-boom jet and engine exhaust with respect to the aircraft wing in supersonic cruise;

FIG. 14 is a sectional view as in FIG. 13, illustrating the relationship between the anti-boom jet, the engine exhaust and the wing during aircraft climb;

FIG. 15 is a diagrammatic view as in FIG. 11, showing an alternate form of anti-boom jet -producing system as arranged on one side of a supersonic transport, utilizing main thrust engines forward and aft.

FIG. 16- is a diagrammatic view as in FIG. 11 showing another alternate form of anti-boom jet-producing system utilizing a separate unit powered by the main thrust-producing engines and as arranged on one side of a supersonic transport;

FIG. 17 is a diagrammatic view as in FIG. 11 showing an alternate removable form of antiboom jetproducing system as arranged on one side of a supersonic transport, utilizing an additional jet engine;

FIG. 18 is a diagrammatic view as in FIG. 11, showing a further alternate form of anti-boom jet-producing system, utilizing a separate unit receiving air ducted from the main thrust-producing engines and as incorporated on one side of a supersonic transport;

FIG. 19 is a diagrammatic view as in FIG. 11, showing another alternate form of anti-boom jet-producing system as arranged on one side and one wing of a supersonic transport;

FIG. 20 is a diagrammatic view showing in greater detail the anti-boom jet-producing means shown in FIG. 11.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT The present invention is concerned with complex and highly sophisticated concepts and devices in the field of aerodynamics and particularly in the field of supersonic flight. As competence in this art requires a considerable level of knowledge, itmay therefore seem to one skilled in the art that the following description is perhaps, in some parts, somewhat superficial. However, it should be appreciated that a detailed description of all of the concepts, qualifications and calculations implicit in the description would require a burdensome amount of time and space while contributing little to an adequate understanding of the invention. Hence the description will cover the essential principles and phenomena necessary to an understanding of the claimed invention while avoiding those related theoretical and engineering aspects within the competence of those skilled in the art. The description will be presented in the following topical order:

Brief Analysis of the Sonic Boom Signature p. l5 Effects of Altering the Boom Signature p. i7 Mechanics of Shock Wave Propagation in Altering the Boom Signature p. 21 Mechanics of Shock Wave Interaction in Weakening the Wing Shock p. 28 Mechanics of Shock Wave Interaction in Maintaining Lift p. 33 Achieving Improved Performance of Supersonic Aircraft p. 38 Means for Producing the Anti-boom Jet p. 40 Use of the General Electric Company's Prototype GE4/J5P Engine in Producing the Anti-Boom Jet p. 48 Arrangement of the Anti-Boom Jet Means on an SST p. 55 Arrangement of the Anti-Boom Jet on the Boeing Company Prototype 2707-300 p. 58

BRIEF ANALYSIS OF THE SONIC BOOM SIGNATURE FIG. 1 illustrates an aircraft 1 flying a supersonic speed and shows the propagation of the major shock waves, that is, the nose shock wave 2, the wing shock wave 3, and the tail, or trailing shock waves 4, to the ground level 5. It will be seen that in the near field, the shocks due to aircraft details are apparent, while in the mid-field region, only the major shocks survive, and in the far field, the signature is frozen" in an asymptotic N -wave 6, which is independent of the aircraft details. The mid-field peak overpressure, A p,,,, will be larger than that at the far field, A p,, since the strength of the shock waves decays during propagation through the atmosphere. The signature measured on the ground will be in the form of a far-field, asymptotic N-wave 7, that is, an N-shaped pressure-vs-time or space profile. The forward portion 7a of the signature is characterized as a sudden rise in pressure A p, above the ambient atmospheric pressure, followed by a gradual decay. This region of overpressure is composed of the leading nose shock wave 2 and the wing shock wave 3, since, during propagation, the wing shock wave travels faster in the heated wake of the nose shock wave and tends to catch up with it as seen in the figure. As previously stated, the

peak overpressure is qualitatively dependent on various aircraft and atmospheric parameters but these may be roughly divided into two interrelated ground level boom components. These two components may be referred to as the volume boom A p and the lift boom A p, The magnitude of the components may be expressed respectively as:

[V h In) where k and k are shape factors of the order of 0.5 to 1.0 and vary slowly with changes in aircraft shape; p and p, are the atmospheric pressure at flight altitude h and at ground level respectively; M is the cruising Mach number of the airplane at altitude h; d and W are the respective characteristic diameter of the airplane volume, and the airplane weight; I is the overall length of the airplane, while I is the characteristic length of the lifting surface.

It will be seen from equations (1a) and (lb) that A 12,, decays as the one-half power of p, at flight altitude h, while A p,, is independent of p Since the atmospheric pressure decays exponentially with increasing altitude, A p decays more rapidly than A p,, as h increases, so that at low altitudes, A p dominates, while at the high cruising altitudes of the heavy commercial supersonic transports, A p,, becomes the larger component. These relationships are plotted approximately in FIG. 2 for a heavy aircraft and taking the interference between the two interrelated components into account. At SST cruising altitudes, therefore, the reduction of the lift boom A P is a major concern. It will also be seen that since increasing scale lengths I and I will reduce the boom according to the A power law, many previously proposed solutions to the boom have gone in that direction and have suggested the creating of a phantom airplane of increased effective length over the actual physical length of the aircraft.

In contrast, the present invention proposes to weaken the lift boom by operating on the wing shock to weaken it and thus achieve a greater separation between it and the nose shock in the far field. Such an approach has generally been avoided in the past for fear of destroying the lift on the aircraft in the process. However, before presenting an explanation of how the wing shock may be weakened while the overall lift from the wing may be maintained, the benefit to be gained from separating the wing shock from the nose shock will be considered.

EFFECTS OF ALTERING THE BOOM SIGNATURE In the series of FIGS. 3 a-e and 4 a-e, two series of boom signatures are schematically shown. FIG. 3a represents a far field asymptotic N-wave 8 formed by a nose shock signature of unit strength, )t,,=l.0, reinforced by a wing shock signature of strength A =1.5. The peak overpressure of the boom signature is 2.5 units. The dotted line indicates the forward portion of the elementary N-wave 8a associated with the leading nose shock. It is formed by the nearly conical shock resulting from the pointed nose of the supersonic aircraft from which it emanates, followed by expansion waves or fans which are generated by the gradual change of contour along the nose region and over the maximum section into the waist of the fuselage ahead of the wing root. The details of the lower magnitude aft portion of the wave 8a are omitted for clarity. The length O-N of the overpressure region of this nose shock signature consists largely of the forward displacement D,, of the actual nose shock from the hypothetical Mach cone passing through the nose of the aircraft at the given instant (FIG.

As previously indicated, an actual shock wave produced by a supersonic aircraft propagates at a speed somewhat greater than the speed of sound, that is, faster than a Mach wave originating from the same point, which propagates at the local speed of sound in the atmosphere. Thus, in propagating to the ground, an actual shock wave gains a distance D on an hypothetical Mach cone originating from the same point. As shown in FIG. 5, the actual nose shock front 9 preceeds the nose Mach cone 9a by a distance D, and the wing shock front 10 preceeds the wing Mach cone 10a by a distance D,,,; the displacement between the nose shock and the wing shock fronts is D The distance, D for the nose shock, for example, as will be more fully explained below, may be of the order of 200 to 300 feet for an aircraft 300 feet in length, flying at 60,000 feet altitude. The zero over-pressure point N of this nose N- wave 8a will be only about feet behind the Mach cone from the nose.

The dash-dot line in FIG. 3a indicates the wing shock-associated lift N-wave 8b in a position almost coinciding with the leading nose shock signature 8a. The zero over-pressure point W of the wing or lift N-wave 8b may be about feet behind the nose Mach cone. Again, much of the length of the overpressure region of the lift N-wave 8b is due to the wing shock displacement D ahead of the wing Mach cone which, as will be presently explained, may be of the order of 300-400 feet for the given example. The resultant boom signature 8 is roughly obtained by the superposition of the two elementary N-waves 8a and 8b, which'gives a peak over-pressure of about 2.5 units. v

Now, suppose that the wing shock signature 8b is displaced toward the rear by l/ 10 of the over-pressure length O-N of the nose signature 8a. The superposition of the wing shock overpressure onto the lower nose shock overpressure will lead to a peak overpressure of 2.4 units as illustrated in FIG. 3b. If the wing shock signature 8b should be displaced rearward by b or even 1 overpressure length O-N as shown in FIGS. 30 and 3d, respectively, the peak overpressure level would be reduced to 2 and 1.5 units, respectively, corresponding to 20 and 40 percent reductions of the boom signature overpressure.

The diagrams in FIGS. 3a-d depict situations wherein the entire wing signature over O-W is rearwardly displaced, while nose shock 0'N remains unchanged. This type of signature movement may be accomplished by shifting the entire wing structure with respect to the nose of the aircraft. However, it should be noted, in these cases, the total positive impulse is not substantially changed, although the peak overpressure is significantly reduced.

Suppose now, the wing signature 8b is displaced rearwardly by weakening and slowing down the wing shock without affecting the zero point of the wing wave system. This is the situation illustrated in the diagrams shown in FIGS. 4a-d. It will be seen thatnot only will the peak overpressure be reduced to the level achieved by the previously mentioned wing structure shift, but also that the positive impulse will be reduced as much as the peak overpressure. This can be accomplished if the shock from the leading edge of the wing is reduced in strength while maintaining the lift fairly constant and without changing the position of the wing trailing edge, so as to maintain the overpressure zero point substantially at W. In actuality, under the constraint of a constant lift from the wing surface, the weakening of the shock from the leading edge of the wing will likely be accompanied by some slight rearward shift of the zero point W of the signature 8b. The effects on the reduction of the peak overpressure and the positive impulse in an actual situation will thus be modified slightly from the extreme situation described in the series of FIGS. 4a-d, toward the other extreme situation described in the series of FIGS. 3a-d.

In any event, the diagrams in FIGS. 3a-d and 4a-d show that the peak overpressure of the boom may be reduced from 2.5 units to 1.5 units when the signature of the wing shock is shifted back to the zero overpressure point N. However, if we ignore the complicated effects on the rear half of the boom signature, the peak overpressure can be lowered even further to the level of the leading nose shock by shifting the wing shock signature further back into the valley behind N. It will be seen that this is possible in view of the fact that the wing shock strength k =1 .5 is larger than the nose shock strength A =l.0. But, if, conversely, the aircraft structure were modified to have a fat nose so that A =1.5, while A =1.0, there would be little advantage in shifting the weaker wing shock signature as far back as point N, since the peak overpressure would remain at the 1.5 units level of the nose shock for any position of the wing shock within the last A; of the nose signature O-N, and behind. This observation is of importance in the event that an extensive redesign or a new design of supersonic aircraft is considered from the particular point of view of reducing the sonic boom.

Another operation in the boom signature will permit the peak overpressure level illustrated in FIGS. 3d and 4d to be reduced below 1.5, even though the wing shock signature is shifted only to the zero overpressure point N of the nose shock. This may be accomplished by slightly strengthening the nose shock from A,,=1.0 to about 1.25 units as shown in FIGS. 3e and 4e and correspondingly shifting the zero overpressure point N upstream to N. This strengthened nose shock signature 8a will then lead the wing shock even further so that the wing shock signature 8b although still of strength A =l .5, will now fall slightly in the valley behind N and rise to a peak overpressure of 1.25 also. Thus, the strengthening and the upstream shift of the nose shock will lead to a reduction of the peak overpressure. This can be achieved, for example, by increasing the nose angle of the aircraft and shifting the maximum sectional area of the nose forward, or as will be explained later, by proper orientation and other adjustments of the anti-boom jet stream.

In order to further explain the intricacies of the principle and method of operation of the present invention, the mechanics of shock wave propagation particularly in connection with altering the boom signature will now be considered in some detail.

MECHANICS OF SHOCK WAVE PROPAGATION IN ALTERING BOOM SIGNATURE The strength of a shock wave has been characterized by A which is a ratio defined as the pressure rise across the wave p-p divided by the air pressure p, ahead of the wave, i.e.,

where p is, of course, the pressure in the region behind the wave. The dynamic laws governing the propagation of a shock wave in a flow field are the conservation laws of mass, momentum and energy. For propagation in the atmosphere, air may be considered as an ideal gas with constant specific heat ratio y=l .40. So, if 8 is the angle of inclination of the air flow relative to the local shock front 11, as shown in FIG. 6, and M is the local flow Mach number, then M, =M sinO will be the Mach number of the speed of propagation of the shock front 11 in the direction normal to the front, and the propagation velocity of the shock front 11 may be expressed in terms of the local speed of sound a, as u, M a where:

It will be seen that for a very weak shock, where )t is very small, the wave propagates at the speed of sound (M,,=l and is thus called a Mach wave. In threedimensional space, such a weak wave frontgenerated by a point source in supersonic motion is in the form of a cone called a Mach cone 12, also shown in FIG. 6. For a wave generated by an aircraft, the velocity of propagation of the shock front 11 can be resolved into a component S, perpendicular to the ground and a horizontal component 8,, parallel to the ground in the direction of cruising flight. For the leading nose shock propagating into an undisturbed quiescent atmosphere with local speed of sound and an aircraft cruising spee u M a we have:

For shock waves propagating into disturbed atmosphere, all the kinematic quantities in these equations, that is, 14,, 0, a, and M should be referred to the local relative air velocity which is not necessarily parallel to the ground. In such event, equations (3a) and (3b) still apply, but the subscripts v and h designate the velocity components perpendicular and parallel to the local flow direction rather than to the ground;

During propagation, a shock wave decays in strength unless suddenly reinforced by another shock wave catching up from behind and while the speed of sound increases in the atmosphere toward the ground level, so does the speed of propagation of a shock wave. Also, since the pressure always increases across a shock wave, A is always positive no matter how small, so that, as will be seen from equation (2), a shock wave will always lead a local Mach wave whose A =0 in its propagation. As a result, a horizontal separation D between the shock front 11 and the associated hypothetical Mach cone 12 occurs, which increases with the'distance of propagation. At a height h below the cruising altitude of an aircraft and immediately underneath the flight path this separation distance D is given by the integral:

w evil" where y is the length of the aircraft, and y represents the vertical distance expressed in terms of y,,. The lower limit of integration y l'represents the reference altitude one aircraft lengthbelow the cruising altitude. The quantities A, 14,, a, and M, inside the integral are all functions of y. 1

When the rate of decay of the shock strength A is estimated by ray acoustics in real atmosphere, a fairly narrow upper and lower bound of the separation distance D in the field far from the aircraft can be obtained to replace the integral in equation (4), thus D= xucyom (5) where the subscript g refers to ground level and subscript refers to the reference altitude. H is the density scale height of the exponential atmosphere, generally taken as about 30,000 feet.

In concrete terms using, for example, values based on those for the previously mentioned Boeing Company SST prototype 2707-300, with y =300 ft. and h, 60,000 ft., a conical nose shock of strength )t 0.1 will have a separation distance D on the ground lying between 260 and 320 feet as computed from equation Note that the upper bound of 320 feet cannot be exceeded even if h, is much greater than H. With h, approximately equal to 2H, both the upper and the lower bound will be about 80 percent of these values, that is, about 200 to 260 feet, which may actually be a more accurate estimate. For a rear or rail shock of similar strength, the signature would lag a similar amount. Thus, the overall length of the boom signature may be 700 to 900 feet or about 2 to 3 lengths of the aircraft.

The wing shock will be stronger than the nose shock and is essentially planar near the wing. lt propagates into an atmosphere which has been set in motion with a velocity u, sin0/[l (7+1 /27) A], by the leading nose shock in a direction normal to the shock front. This disturbed air is at a slightly higher pressure and temperature than the ambient air into which the nose shock is propagating. From FIG. 5, it can be seen'that the motion of the wing shock wave relative to the moving nose shock wave 9 will hence consist of two displacements:

i. The displacement D of the wing shock front 10 from the wing Mach cone 10a, which displacement may be evaluated from equation (4) or (5) and (ii) the displacement D of the nose shock wave 9 from the wing Mach cone 10a. Then, D-, D, D D D or D D, (D D where D. is the distance between the Mach cones from the aircraft nose and the wing. Thus, the displacement D between the nose shock 9 and the wing shock 10 is a function of the difference between D and D The wing is generally located in the waist of the fuselage in accordance with the well-known aerodynamic Area Rule behind the dividing Mach line of the fuselage, which will be explained more fully in connection with FIG. 7a. The Mach cones from this region are retarded by the overexpanded flow behind the nose Mach cone and will lag behind the undisturbed Mach cone from the nose with increasing horizontal separation between them, so that D, will be somewhat larger than the distance D, between the aircraft nose and wing. Accordingly, the difference between the values of D of the wing shock and of D of the nose shock evaluated from equations (4) or (5), i.e., D WN will slightly overestimate the decrease of the horizontal separation distance between the two shocks from the initial separation.

The strength of the wing shock varies along the wing span and approaches zero at the wing tip. Thus, although the wing tip is initially nearest to the nose conical shock, the wing shock from the tip will not catch up to the nose shock. The wing shock strength increases inboard toward the fuselage and despite the larger initial separation, the wing shock inboard, if sufficiently strong, will catch up and reinforce the nose shock. For example, the major lifting shock from the wing in the prototype 2707-300 will be initially 100 to 150 feet behind the nose. If the reference strength A, of the wing shock is, say, about 0.05 larger than that of the nose shock, the signature of the wing shock will reinforce that of the nose shock and the asymptotic N-wave boom signature will result. If it is appreciably lower, however, the wing shock will not have caught up by the time the shock system reaches the ground. The boom signature at the ground level will then contain multishocks as in the mid-field profile rather than the asymptotic far field form.

It will be appreciated by those skilled in the art that the foregoing presentation introduces theshock expansion type argument intothe quasi-linear theory of G. B. Whitham, which is particularly set forth in the articles The Flow Pattern of a Supersonic Projectile, Communication in Pure and Applied Mathematics, Vol. 5, 1952, pp. 301-348, and On the Propagation of Weak Shock Waves," Jour. of Fluid Mechanics, Vol. 1', 1955, pp. 290-318. The foregoing modification is in recognition of the fact that shock waves do originate directly from certain surfaces of the aircraft rather than emerge slowly from the congruence of isentropic compression prior to their annihilation or merging with other shocks. In this sense, ray acoustics can be used in estimating the decay of individual weak shock waves with some modification. The presence of an absolute upper bound of D for infinite distance of propagation and the possibility of two shocks remaining separated in real atmosphere if the upper bound of D D- is less than their initial separation are in agreement with W. D. Hayes predictions of the aging of signature shape, as explained in the article Self-similar Strong Shock in an Exponential Medium, .lour. of Fluid Mechanics, Vol. 32, 1968, pp. 305-316.

-The results of optimization studies based on Whithams theory by presuming the far field, asymptotic N-wave signature, are very much different from the results obtained when mid-field characteristics of the signature are taken into consideration. Based on the far field results, none of the many boom reduction schemes that have been proposed are practical and optimization studies for redistributing the lift suggest the desirability of heavier loading near the wing leading edge to reduce the boom. However, the development in the foregoing presentation suggests, on the contrary,

The extension of the wing structure or the rearward shift of the whole wing would serve to appropriately vary the wing shock signature. However, although desirable, such a modification could presently only be considered in the design of a new aircraft as the existing SST prototypes would not be readily adaptable. Such a modification would also be likely to involve an extensive fixed structural weight penalty on the aircraft during flight periods when boom reduction is not needed. Accordingly, a preferred adaptation of existing SST aircraft will be described for achieving the weakening of the wing shock which does not involve extensive modiflcation of the wing structure.

MECHANICS OF SHOCK WAVE INTERACTION IN WEAKENING THEWING SHOCK We shall now turn to the explanation of how the wing shock is weakened. For illustrative purposes, as shown in FIG. 7a, a bi-convex airfoil section at an angle of attach a less than the nose angle, is taken as a representative two-dimensional section across the wing of an aircraft in supersonic cruise. A pair of shock waves 21 and 22 are created at the leading edge of the airfoil 20, one on the upper and one on the lower surface. The wave 22 on the underside of the wing edge will be stronger and propagates toward the ground. The biconvex contour of either the upper or the lower wing surface generates a family of expansion waves 23 distributed over the entire section 20, with the pressure over the rear portion being considerably below that over the upstream portion of the wing section. This family of expansion waves on eitherside can be divided into two groups separated by a dividing Mach line 24. The waves 23 in front of the dividing Mach line 24 on the lower side will eventually catch up to the lower leading edge shock 22 to account for part of the decay of the shock wave during its propagation to the ground. The expansion waves 23 behind the dividing Mach line 24 overexpand the air and will not catch up to the shock wave 22 from the leading edge but will eventually be merged into the shock wave 25 from the trailing edge. The trailing shocks 25 are formed when the two air streams 26 and 27, moving at different velocities along the upper and the lower wing surfaces, respectively, meet at the trailing edge 28. The trailing shocks 25 act to turn the two streams 26 and 27 to again flow in the same direction and at the same pressure aft of the wing.

A diagram of the static pressure on the upper and the lower wing surfaces is given in FIG. 7b. The lift on the wing is the difference between the component of force acting on the upper surface P and that acting on the lower surface P, of the wing in the direction of the lift. It will be seen in FIG. 7c that the lift is distributed fairly uniformly over the wing chord and is due to the excess overpressure on the forward half of the lower (or pressure) surface and to the excess underpressure on the aftward half of the upper (or suction) surface. The center of pressure is nearly the midchord. Such pressure variation over a wing surface in a supersonic flow stream is produced by the shock and expansion waves so that certain basic properties associated with a shock wave should be kept in mind, particularly preparatory to the subsequent discussion. Firstly, pressure, density and temperature increase across a shock wave while the velocity of flow decreases. The direction of flow is deflected toward the shock wave and the deflection is a function of the pressure change across the wave and the Mach number of the flow. The deflection decreases with increasing Mach number and the pressure ratio across a shock wave increases with increasing Mach number. Weaker shock waves propagate more slowly than stronger waves and the converse rarefaction or expansion waves will produce results directly opposite to that of shock waves.

Now, as shown in FIG. 8a, a shock wave S from a wing leading edge 30 deflects the ambient air flow through an angle 6,. This is the angle through which the air flow next to the aircraft surface must turn to accommodate the configuration of the aircraft wing. It is this deflection angle 6 which determines the strength of the shock waves originating from the various surfaces 0 the aircraft. It is given by the relation: t

the deflection angle 8 for a given value of flow Mach number M as in FIG. 8b, is known as the 3-8 shock polar. As shown in FIG. 8b, the origin, or zero point of the plot designates the zero deflection angle, i.e., 8=0, and the air pressure p, in front of a shock wave. Each curve through the point 0 is a shock polar for flow of a given Mach number, M. Each point on the curve 0-1 for example, represents the pressure p behind a shock wave after it has deflected the air stream through an angle 6.

Consider now that there is present in the atmosphere a contact discontinuity, that is, a surface separation two parallel air streams which have the same pressure but which are moving with different Mach numbers. (They may also differ inmany other ways). As shown in FIG. 811, when a shock wave S is incident on a contact discontinuity 31, it is partially transmitted arid refracted, and partially reflected. This interaction is conveniently analyzed with the p-6 shock polar in FIG. 8b, taken together with FIG. 8a. The curve S represents the shock in upper stream 32 of Mach number M, in FIG. 8a, which raises the stream pressure from p in region to p in region 1 and deflects the stream 32 through angle 8,. The shock S is incident on the contact discontinuity 31 at point A. Now, if the shock S were transmitted unchanged in strength into lower stream 33, whose Mach number is M, which is less than M, it would cause the shock transition in the lower stream 33 to go from the pressure at point 0 on the plot to that at point 2" and deflect the stream 33 through an angle 8 larger than 8,. The two neighboring streams 32 and 33 would then diverge. On the other hand, if the shock S were transmitted so as to deflect the lower stream 33 through the same angle 8,, the pressure in the stream 33 would be raised only to the point 1, on the shock polar, lower than that of point 1. Therefore, the condition of equal pressure between the two neighboring parallel streams 32 and 33 can only be achieved by the creation of an expansion of rarefaction wave R reflected from the point A. The isentropic expansion from the Mach number at' point 1 follows the line of isentrope 1-2 in the p-8 polar diagram. The intersection of this isentrope with the shock polar 0-1'-2 at point 2 or 2 determines the state of the region 2 behind the expansion fan R in the upper stream 32 and the state of region=2 behind the transmitted shock SE in the lower stream 33. 2 and 2' are the same point in the p-6 polar diagram since the two states are at the same pressure and deflection angle although they differ in Mach numbers and many other quantities. We see then that the transmitted shock S into stream 33 of lower Mach number is weaker than the incident shock S in stream 32 with a higher Mach number.

FIGS. 8c and 8d demonstrate the situation for a stream 34 whose Mach number, M, is greater than M. A shock wave C will be reflected from point A and the transmitted shock into a stream 34 of higher Mach number is stronger than the incident shock S When the difference between the two Mach numbers M and M is small, the reflected shock wave C is weak and may be treated as an isentropic compression wave. The

isentrope 1-2 is defined by the Prandtl-Meyer'relation:

By eliminating 8 and M at the-state 1 between equation (7) and the differential relation of equation (6), a complicated nonlinear differential equation results. This relation gives the variation of the transmitted shock strength crossing a contact discontinuity of small Mach number difference. The nonlinearity of this relation is very important as will be appreciated by those skilled in the art.

Turning next to FIG. 9a, consider the crossing of a shock wave S through three parallel streams. Stream 36 is of higher Mach number, stream 37 is of lower Mach number, and stream 38 is of equalMach number with the upper stream 35. Curve 0-! in the P-6 polar of FIG. 9b represents the shock transition of the upper stream 35with Mach number M from state 0 to state 1. A shock C is reflected from point A on the contact discontinuity 39. The reflected shock C, brings the stream 35 to state 2 in balance with state 2' in the stream'36 of higher Mach number M, which results from the crossing of the transmitted shock S This transmitted shock S is incident on the lower contact discontinuity 40 at point B and is refracted into the next stream 37 of Mach number M which is lower than the Mach number M. An expansion wave R, will be reflected from B bringing the stream 36 to the state 3 in balance with the state 3E brought about by the transmitted shock S in the lower Mach number stream 37. The shock is then incident on the lowest contact discontinuity 41 at C and emerges into the stream 38 whose Mach number M" is equal to M. A shock C3345 is reflected from point C to bring the state 4B in balance with the state 4" in the stream 38 brought about by the emerged shock D Since the Mach number of the flow at 2' is higher than the Mach number at 2 and the Mach number of the flow at 3 is even higher than that at state 3E, the p-& polar diagram in FIG. 9b can be used to determine the relative positions of these reflected shock curves and the various states. It will be seen that the pressure rise across the emerged shock S (point 4") is considerably less than the pressure rise across the original shock S (point 1). Therefore, the emerged shock strength is considerably weakened by the two intervening streams.

It should be understood-that the situations described in FIGS. 8a-d and 9a-b are idealized in that each stream is of uniform Mach number and the contact discontinuities remain sharp. In actuality, the sharp discontinuities will be diffused by molecular and turbulent fluctuations and will be replaced by mixing zones of gradual change of various quantities from one stream to another. Isentropic compression waves will replace the reflected shock waves C and C m; from A and C. Each set of compression waves will converge to form a single shock eventually if not intercepted. The illustrations in FIGS. 8ad and 9a-b, also describe the situation when the thickness of the mixing zone is considerably less than the width of each stream. However, the p-8 polar diagrams show that such modification is not of fundamentalimportancei l I Now, if the shock S represents a shock emanating from w wing leading edge with thelower wing surface serving as the solid boundary, it will be seen that an interaction will jet streams of different Mach numbers weakens the emerged shock S This weaker shock upon propagating to the ground will produce a smaller pressure rise and will lag behind the signature of the original wind shock-S that is, the shock not intercepted by the jet streams.

MECHANICS OF SHOCK WAVE INTERACTION IN MAINTAINING LIFT It must be emphasized that although the original wing shock is weakened, the lift force on the wind is not 'adversely effected but indeed may be enhanced by the application of such jet streams. This feature will now be explained with reference to FIGS. 10a and 105. An artti-boom jet stream 50 generated parallel to ambient air stream 51 and at the same pressure but higher Mach number and at a vertical distance it, below the undersurtace of a bi-convex type Wing 52, is shown in FIG. 10a. When shock wave S from the leading .edge of wing 52 encounters the upper boundary 53 of this jet stream as a contact discontinuity, an interaction of the type describedin connection with FIG. 8c results and a compression wave C is reflected into the deflected 

1. Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of shock waves reflected from the interaction will strike the adjacent side of the surface aft of its dividing Mach line, thereby increasing the lifting force thereon.
 2. Method as in claim 1, including the step of reducing the angle of attach of said lifting surface in response to the increased lifting force thereon.
 3. Method of reducing the sonic boom from a supersonic aircraft, comprising the steps of: a. creating a contact discontinuity below a major lifting surface of the aircraft by producing a fluid stream having a Mach number greater than the Mach number of its ambient supersonic flow; and b. directing said fluid stream at a region aft and below the leading edge of said major lifting surface, to cause the major lift shock wave propagating therefrom to propagate into and interact with the contact discontinuity beneath the lifting surface such that compression waves resulting from the interaction will strike the underside of the lifting surface, thereby weakening said shock wave and altering the boom signature while maintaining lift.
 4. Method as in claim 3, wherein said major lifting surface is an aircraft wing and said contact discontinuity is created below the wing leading edge a distance of approximately 1/4 of the wing chord.
 5. Method as in claim 3, including directing the fluid stream such that the compression waves reflected from the interaction will strike the underside of the major lifting surface aft of its dividing Mach line.
 6. Method as in claim 5, including directing said fluid stream such that shock waves reflected from the interaction of the contact discontinuity and said shock wave from the wing leading edge strike the aft portion of the underside of the wing, increasing the lift force thereon; and reducing the angle of attach of the lifting surface in response to the increased lifting force produced by the compression waves to further weaken the wing shock wave.
 7. Method as in claim 3, including producing a second fluid stream having a Mach number lower than that of its ambient supersonic flow and directing said second stream toward said shock wave beneath said lifting surface.
 8. Method as in claim 2, including directing said greater Mach number stream between said lifting surface and said lower Mach number stream.
 9. Method as in claim 8, wherein the greater Mach number stream includes air from a source of liquid air.
 10. Method as in claim 8, wherein said greater Mach number stream contains a cooling agent selected from the group consisting of NH3, CO2 and H2O.
 11. Method as in claim 3, wherein said fluid stream is produced by c. capturing ambient air; d. increasing the stagnation pressure of the captured air; and releasing said air as a jet of fluid.
 12. Method as in claim 11 wherein said air is released from a location adjacent and downstream from the aircraft nose, and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
 13. Method as in claim 12, wherein said air is released at a pressure above the static pressure of the surrounding flow to strengthen the nose shock wave, thereby altering the boom signature.
 14. Method as in claim 12, wherein said air is released at a pressure below the static pressure of the surrounding flow to weaken the nose shock wave, thereby altering the boom signature.
 15. Method of improving the performance of supersonic aircraft, comprising the steps of: a. creating a contact discontinuity adjacent a lifting surface of the aircraft by producing a fluid stream having a Mach number lower than that of its ambient supersonic flow; and b. directing the stream such that the contact discontinuity interacts from upstream with a leading shock wave from said lifting surface in a region from which a majority of expansion waves reflected from the interaction will strike the adjacent side of the surface forward of its dividing Mach line, thereby decreasing the drag force thereon.
 16. Method of decreasing the overpressure and positive impulse of an aircraft sonic boom signature, comprising creating a contact discontinuity in the near field of the aircraft for preventing the leading wing shock wave from catching the leading nose shock during propagation thereby shifting the wing shock signature and the nose shock signature with respect to each other in the far field, so that the peak overpressure of the wing shock signature will rise in the valley of the nose shoCk signature at ground level
 17. Method as in claim 16, wherein the wing shock signature is shifted by weakening the wing shock to decrease its propagation velocity.
 18. Method as in claim 16, wherein the nose shock signature is shifted by strengthening the nose shock to increase its propagation velocity.
 19. Means for improving the performance of a supersonic aircraft comprising: a. means for producing a fluid stream having Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with a shock wave propagating adjacent a lifting surface of the aircraft in such manner that shock waves reflected from the interaction will strike the adjacent side of the lifting surface aft of its dividing Mach line.
 20. Means as in claim 19, wherein the means for producing the fluid stream includes a source of liquid air.
 21. Means as in claim 19 wherein the means for producing the fluid stream comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release into said fluid stream.
 22. Means as in claim 19, wherein the lifting surface is an aircraft wing and said fluid stream producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
 23. Means as in claim 19, wherein the means for producing the fluid stream comprises a main thrust-producing engine of the aircraft.
 24. Means as in claim 23, wherein the means for producing the fluid stream comprises: i. a source of liquid air; ii. means for conducting liquid air from said source to cool said main thrust-producing engine; iii. means for receiving said air after cooling said thrust-producing engine and collecting it at a high stagnation pressure; and iv. means for expanding and releasing said high stagnation pressure air into said fluid stream.
 25. Means for reducing the sonic boom from a supersonic aircraft comprising: a. means for producing a fluid stream having a Mach number greater than that of its ambient supersonic flow; b. means for mounting said fluid stream-producing means on the aircraft; and c. means on the aircraft for directing said fluid stream to interact from upstream with the leading shock wave propagating beneath a major lifting surface of the aircraft, in a region aft and below the leading edge of said lifting surface from which the compression waves resulting from the interaction will strike the underside of the lifting surface, to weaken and decrease the propagational velocity of the shock wave, thereby altering the aircraft boom signature while maintaining the lifting force on the lifting surface.
 26. Means as in claim 25, wherein the means for producing the greater Mach number stream comprises a main thrust-producing engine of the aircraft.
 27. Means as in claim 25, wherein the means for producing the fluid stream comprises a turbofan engine.
 28. Means as in claim 25, wherein the lifting surface is an aftwardly positioned aircraft wing and said fluid stream producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
 29. Means as in claim 25, including: d. means on the aircraft for producing a fluid stream having a Mach number lower than that of its ambient supersonic flow; and e. means for directing said lower Mach number stream immediately below the greater Mach number stream.
 30. Means as in claim 29, wherein the means for producing the low Mach number fluid stream comprises a main thrust-producing engine of the aircraft.
 31. Means as in claim 25, wherein tHe means for producing the greater Mach number stream comprises a source of liquid air.
 32. Means as in claim 31, comprising means for conducting said liquid air from said source to cool a main thrust-producing engine of said aircraft.
 33. Means as in claim 25, wherein the means for producing the fluid stream comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release as said fluid stream.
 34. Means as in claim 33, comprising means for introducing a cooling agent into said stagnation pressure-increasing means.
 35. Means as in claim 25, wherein the means for producing the fluid stream comprises a turbojet engine.
 36. Means as in claim 35, wherein said turbojet engine comprises: i. a source of liquid air; ii. means for conducting liquid air from said source to cool said turbojet engine; iii. means for receiving said air after cooling said turbojet engine and collecting it at a high stagnation pressure; and iv. means for expanding and releasing said high stagnation pressure air into said fluid stream.
 37. Means as in claim 36, wherein said releasing means is located above the exhaust stream of said turbojet engine.
 38. Supersonic aircraft comprising: a. means mounted on the aircraft for producing a jet of fluid with a higher Mach number than that of its ambient supersonic flow; b. lift-producing means; and c. means directing said fluid jet under said lift-producing means to interact from upstream with the lift shock wave beneath the lift-producing means.
 39. Aircraft as in claim 38, wherein the means for producing the fluid jet is a turbofan engine.
 40. Aircraft as in claim 38, wherein the lift-producing means is the aircraft wing and said jet-producing means is mounted adjacent and downstream from the aircraft nose and upstream of the lower nose contour dividing Mach line to alter the nose shock strength.
 41. Aircraft as in claim 38, wherein the means for producing said fluid jet comprises a source of liquid air.
 42. Aircraft as in claim 41, wherein said higher Mach number jet-producing means comprises a main thrust-producing engine of said aircraft and means for conducting said liquid air from said source to cool said main thrust-producing engine.
 43. Aircraft as in claim 38, comprising means mounted on the aircraft for producing a jet of fluid with a lower Mach number than that of its ambient supersonic flow.
 44. Aircraft as in claim 43, wherein said lower Mach number jet-producing means comprises a turbojet engine mounted forward of said lift shock wave, the exhaust from while engine comprises a low Mach number jet and interacts with said lift shock wave.
 45. Aircraft as in claim 44, wherein said higher Mach number jet-producing means comprises said turbojet engine.
 46. Aircraft as in claim 38, wherein the means for producing the fluid jet comprises: d. inlet means for taking in ambient air; e. means for increasing the stagnation pressure of the air taken in; and f. nozzle means for expanding the increased stagnation pressure air for release into said fluid jet.
 47. Aircraft as in claim 46, comprising means for introducing a cooling agent into said stagnation pressure-increasing means.
 48. Aircraft as in claim 46, wherein said stagnation pressure-increasing means comprises a compressor.
 49. Aircraft as in claim 48, wherein said compressor is powered by a main thrust-producing engine of the aircraft.
 50. Aircraft as in claim 48, comprising means for conducting ambient air taken in by a main thrust-producing engine of said aircraft to said compressor. 